脉冲燃烧风洞与常规高超声速风洞数据相关性研究  

Analysis of data correlation between combustion heated impulse facility and hypersonic wind tunnel

作　　者：贺元元[1] 吴颖川[1] 张小庆[1] 林其[1] He Yuanyuan;Wu Yingchuan;Zhang Xiaoqing;Lin Qi(Science and Technology on Scramjet Laboratory,Hypervelocity Aerodynamics Institute of China Aerodynamics Research and Development Center,Mianyang Sichuan 621000,China)

机构地区：[1]中国空气动力研究与发展中心超高速空气动力研究所高超声速冲压发动机技术重点实验室,四川绵阳621000

出　　处：《实验流体力学》2018年第3期64-68,共5页Journal of Experiments in Fluid Mechanics

摘　　要：不同风洞因模拟来流参数不同,对高超声速飞行器气动力试验结果影响很大。总结了脉冲燃烧风洞和常规高超声速风洞不通气标模的试验和计算结果,分析了水凝结、雷诺数、壁温比对模型气动性能的影响规律。脉冲燃烧风洞获得的气动性能变化规律与常规高超声速风洞一致,脉冲燃烧风洞获得的阻力系数比常规高超声速风洞阻力系数大15%左右,其中雷诺数影响较小,在5%以内,壁温比影响较大,在10%以上。结合数值计算对造成差异的原因进行分析,认为壁面传热对边界层速度型的影响是主要因素。The simulation parameters of different hypersonic test facilities have large influence to the test results.A typical lift body aerodynamic test was conducted in the combustion heated impulse facility and hypersonic wind tunnel.The influence of Reynolds number and wall temperature ratio was analyzed.The aerodynamic changes of the combustion heated impulse facility test were in consistence with those of the hypersonic wind tunnel test,but its drag coefficients were about 15%larger.Numerical analysis indicates that the differences caused by Reynolds number were about 5%,and those caused by the wall temperature ratio were about 10%,and therefore the change of velocity boundary layer is the main factor.

关 键 词：脉冲燃烧风洞 常规高超声速风洞 数据相关性 雷诺数 壁温比 

分 类 号：V211.73[航空宇航科学与技术—航空宇航推进理论与工程]